1. Field of the Invention
The invention relates to composite turbine nozzles.
2. Description of Related Art
A typical gas turbine engine of the turbofan type generally includes a forward fan and a booster or low pressure compressor, a middle core engine, and a low pressure turbine which powers the fan and booster or low pressure compressor. The core engine includes a high pressure compressor, a combustor and a high pressure turbine in a serial flow relationship. The high pressure compressor and high pressure turbine of the core engine are connected by a high pressure shaft. High pressure air from the high pressure compressor is mixed with fuel in the combustor and ignited to form a high energy gas stream. The gas stream flows through the high pressure turbine, rotatably driving it and the high pressure shaft which, in turn, rotatably drives the high pressure compressor.
The gas stream leaving the high pressure turbine is expanded through a second or low pressure turbine. The low pressure turbine rotatably drives the fan and booster compressor via a low pressure shaft. The low pressure shaft extends through the high pressure rotor. Most of the thrust produced is generated by the fan. Marine or industrial gas turbine engines have low pressure turbines which power generators, ship propellers, pumps and other devices while turboprops engines use low pressure turbines to power propellers usually through a gearbox.
The high pressure turbine has a turbine nozzle which is usually segmented and includes an annular ring or row of turbine nozzle segments. Each segment includes circumferentially spaced apart airfoils radially extending between radially inner and outer bands. The airfoils are usually hollow having an outer wall that is cooled with cooling air from the compressor. Hot gases flowing over the cooled turbine vane outer wall produces flow and thermal boundary layers along outer surfaces of the vane outer wall and end wall surfaces of the inner and outer bands over which the hot gases pass.
Conventional segments have been made of metal and, more recently, designs have been developed incorporating composite material for parts of or all of the segments. Composite materials offer greater thermal protection and allow the turbine to operate at higher temperatures and reduce the amount of cooling air. These lead to improved component life and engine efficiency. One particular material is ceramic matrix composites also referred to as CMC.
The composite materials are made of plies. Composite material nozzles and composite articles in general are somewhat brittle as compared to their metallic counterparts. First stage high pressure turbine nozzle segments are frequently simply supported at both the inner and outer ends. The two supports move relative to each other due to changing engine conditions. This makes contact between the nozzle and the supports. The nozzle segment must be able to rock and slide relative to the supports in order to remain in contact with the supports. If uniform contact is not maintained then concentrated stresses will be induced which may result in a part life reduction. Stresses are particularly detrimental to composite material components of the turbine nozzle and its segments. A high pressure drop exists across contact areas between the nozzle and the supports during engine operation making it difficult to maintain contact. The high pressure drop results in increased leakage if uniform contact is not maintained and the increased leakage results in a performance decrease.
The nozzle bands are arcuate to conform to the annulus of the engine. An arc shaped contact between the band and support creates an uneven load distribution as the nozzle rocks and opens large leakage areas. At the inner band especially, available space is small and axial room is limited by adjacent turbine blade attachment. Metal turbine nozzles include a support rail typically cast as part of the nozzle segment, approximately ¾ of the axial distance back from the leading edge of the inner band or platform. This is machined to form a chordal hinge or chordal seal so that geometry local to the contact area is shaped such that the contact occurs along a straight line (chord). This allows the vane segment to rock without changing the load distribution or increasing leakage.
A chordal hinge between the HPT nozzle segment outer band and shroud allows relative rocking without opening excessive leakage paths. The chordal geometry causes the radial distance between the flowpath and load reaction point to vary. The longer distance at the sides of the outer band of the segment can cause bending moments to exceed material allowable for composite turbine nozzle segments. It is desirable to provide a chordal hinge and seal for the outer band of a composite nozzle segment with sufficient strength to prevent bending moments from exceeding material allowable for composite turbine nozzle segments.